A continuing interest exists for improvements in rocket engines, and more particularly for designs that would provide a significant increase in propulsive power, as often characterized by the benchmark of specific impulse, especially as might be compared to conventional chemically fueled rocket engines. Such new rocket engines might be useful in a variety of applications. Launch operational costs might be substantially reduced on a per pound of payload basis, by adoption of a new nuclear thermal propulsion rocket engine design that provides significant improvements in the specific impulse, as compared to existing prior art rocket engine designs. Further, from the point of view of overall mission costs, since the mass of most components of rocket vehicles are proportional to the mass of the propellant, it would be desirable to develop a new rocket engine design that reduces the mass of consumable components necessary for initiating lift off and acceleration to orbital velocity. Such an improvement would have a major impact on the entire field of rocket science from a launch weight to payload ratio basis. Such an efficiency improvement would also facilitate the inclusion of wings and recovery systems that would enable an economic fully reusable launch system with airliner type operations, for example, as described in U.S. Pat. No. 4,802,639, issued Feb. 7, 1989 to Richard Hardy et al., entitled HORIZONTAL TAKEOFF TRANSATMOSPHERIC LAUNCH SYSTEM, the disclosure of which is incorporated herein in its entirety by this reference. And, for missions beyond earth orbit, it would be advantageous, from the point of view of mission duration, to provide a new rocket engine design that reduces not only the payload to launch weight, but also the transit time to the mission objective, by providing high specific impulse, so as to minimize fuel required to achieve high vehicle velocities necessary to accomplish a selected interplanetary mission in a specific time frame. And, it would be desirable to provide such an improved rocket engine that includes components which have been reused and identified as comparatively reliable and cost effective, and thus, minimizes design risk and thus minimizes the extent of testing that may be necessary, as compared to many alternate designs which are subject to stress and strain from temperature and pressure in rocket engine services. Thus, it can be appreciated that it would be advantageous to provide a new, high efficiency rocket engine design which provides a high specific impulse, thus minimizing the launch weight to payload ratio.
In general, the efficiency of a rocket engine may be evaluated by the effective use of the consumable propellant, i.e. the amount of impulse produced per mass of propellant, which is itself proportional to the velocity of the gases leaving the rocket engine nozzle. In nuclear thermal rocket engine systems, the specific impulse increases as the square root of the temperature, and inversely as the square root of the molecular mass of the gases leaving the rocket engine nozzle. Consequently, in the design of a nuclear thermal rocket engine, efficiency is maximized by using the highest temperature available, given materials design constraints, and by utilizing a fluid that has a very low molecular mass for generation of thrust.
A variety of fission based rocket engines have been contemplated, and some have been tested. An overview of the current status of such efforts, and suggestions as to suitable configurations for various missions, was published on Oct. 16, 2014, at the Angelo State University Physics Colloquium in San Angelo, N. Mex., by Michael G. Houts, Ph.D, of the NASA Marshall Space Flight Center, Huntsville, Ala., in his presentation entitled Space Nuclear Power and Propulsion; available at website: ntrs.nasa.gov/search.jsp?R=20140016814. As he notes, the Rover/NERVA program (1955-1973) tested a fission rocket engine design. Further, the most powerful nuclear rocket engine that has been tested, to date, was the Phoebus 2a, which utilized a reactor that was operated at a power level of more than 4.0 million kilowatts, during 12 minutes of a 32 minute test firing. However, it is clear that the various nuclear fission rocket engine designs currently available have various drawbacks, such as excessive gamma radiation production of retained core components, which requires extensive and heavy shielding, if used on manned missions.
One of the more interesting disclosures of a fission based rocket engine was provided in U.S. Pat. No. 6,876,714 B2, issued on Apr. 5, 2005 to Carlo Rubbia, entitled DEVICE FOR HEATING GAS FROM A THIN LAYER OF NUCLEAR FUEL, AND SPACE ENGINE INCORPORATING SUCH DEVICE, the disclosure of which is incorporated herein in its entirety by this reference. That patent discloses the heating of hydrogen gas by fission fragments emitted from a thin film of fissile material, such as Americium metal or a compound thereof, which is deposited on an inner wall of a cooled chamber. However, that device generally describes the use of fissile material in critical mass conditions, and although it mentions the contemplation of sub-critical mass fission arrangements, details of such a condition are scant, if indeed present at all in the description thereof.
Thus, there remains a need to provide a design for a high specific impulse nuclear thermal propulsion rocket engine that simultaneously resolves two or more of the various practical problems, including (a) minimizing the weight of consumables (such as chemical fuel constituents) on a per payload pound basis; (b) avoiding excessive radiation shielding requirements when the design is used in manned missions, by avoiding use of retained radioactive hardware that generates large gamma ray emissions; (c) providing for power control, especially as related to power generation amounts at any given time, by providing for throttling of the fission reaction; and (d) providing a high specific impulse, as compared to prior art rocket engines for space vehicles.